Jet engine cold air cooling system

ABSTRACT

Methods and devices for cooling systems ( 100, 700 ) are provided that are in fluid communication with bleed air from a jet engine compressor. The cooling systems include: a first precooler ( 210 ) receiving bleed air from the jet engine compressor; a heat exchanger ( 730 ) downstream from the first precooler ( 210 ); a cooling system compressor ( 220 ) downstream from the first precooler ( 210 ), wherein the heat exchanger ( 730 ) and the cooling system compressor ( 220 ) are in separate flow paths from the first precooler ( 210 ); a cooling system precooler ( 230 ) downstream from the cooling system compressor ( 220 ); a cooling system turbine ( 240 ) with variable guide vanes—VGT—and downstream from the cooling system precooler ( 230 ); and a discharge conduit ( 245 ) downstream from the cooling system turbine ( 240 ) and the heat exchanger ( 730 ). A bypass line ( 290 ) can also be included that bypasses the cooling system turbine ( 240 ).

PRIORITY INFORMATION

The present application is a national stage application under 35 U.S.C.§ 371(c) of prior filed PCT application serial number PCT/US15/38528,filed on Jun. 30, 2015, which claims priority to U.S. Provisional PatentApplication Ser. No. 62/020,512 titled “Jet Engine Cold Air CoolingSystem” of Leamy, et al. filed on Jul. 3, 2014, and to U.S. ProvisionalPatent Application Ser. No. 62/022,364 titled “Jet Engine Cold AirCooling System” of Leamy, et al. filed on Jul. 9, 2014. The disclosuresof the above-listed applications are incorporated by reference herein.

FIELD OF THE INVENTION

The present disclosure is directed to jet engines and, morespecifically, to utilization of jet engine bleed air for cooling thermalloads associated with the engine or vehicle to which the engine iscoupled.

BACKGROUND OF THE INVENTION

Modern day jet airplanes direct regulated airflow from the jet engine tothe occupied cabins and other areas of the aircraft. This airflow,commonly referred to as bleed air, may be withdrawn from the highpressure compressor (HPC) section of a jet engine. U.S. Pat. Nos.5,137,230 and 5,125,597 describe conventional structures and methodsutilized to direct bleed air into environmental control systems (ECS) ofthe aircraft that further process the bleed air prior to cabinintroduction. ECS incorporate various pieces of equipment such as aircycle machines (ACMs), regulating valves, heat exchangers, and otherapparatus to condition engine bleed air prior to cabin introduction.

Bleed air is commonly extracted from multiple locations along the HPCsection using regulated flow to control the extent to which bleed air iswithdrawn. Among the regulating structures are check valves that operateto allow or discontinue airflow and downstream regulator valves thatreduce the pressure of the withdrawn bleed air before it reaches theECS. This reduced pressure bleed air may be directed to a turbine, wherework is extracted, with the bleed air outlet pressure and temperaturefrom the turbine being significantly reduced. This reduced pressurebleed air remains relatively hot and is thereafter cooled by fan air ina heat exchanger associated with the jet engine conventionally referredto as a precooler. Cooled bleed air output from the precooler isdelivered to the ECS where it may be further cooled and pressuresfurther regulated prior to introduction to the occupied cabins or otherareas of the aircraft. In addition to supplying bleed air to the ECS,the jet engine provides a heat sink that provides precooled air to theaircraft and receives high temperature air from the aircraft in returnas part of a cycle.

Regardless of the structures or methods utilized, one constant hasremained with respect to the bleed air supplied to the ECS: it could beno lower in temperature than the lowest temperature air flowing throughthe jet engine. Moreover, the bleed air has always been regulated fromthe HPC using flow control valves that restrict airflow and areoperative to step down the bleed air pressure prior to reaching theprecooler. Consequently, there is a need in the art for structures andmethods of delivering bleed air to an ECS at temperatures lower than thelowest temperature air otherwise flowing through the jet engine.

BRIEF DESCRIPTION OF THE INVENTION

Aspects and advantages of the invention will be set forth in part in thefollowing description, or may be obvious from the description, or may belearned through practice of the invention.

Cooling systems are generally provided that are in fluid communicationwith bleed air from a jet engine compressor.

In one embodiment, the cooling system includes: a first precooler influid communication with the bleed air from the jet engine compressor; aheat exchanger in fluid communication with and downstream from the firstprecooler; a cooling system compressor in fluid communication with anddownstream from the first precooler, wherein the heat exchanger and thecooling system compressor are in separate flow paths from the firstprecooler; a cooling system precooler in fluid communication with anddownstream from the cooling system compressor; a cooling system turbinein fluid communication with and downstream from the cooling systemprecooler; and a discharge conduit downstream from the cooling systemturbine and the heat exchanger.

In another embodiment, the cooling system includes: a first precooler influid communication with the bleed air from the jet engine compressor; acooling system compressor in fluid communication with and downstreamfrom the first precooler; a cooling system precooler in fluidcommunication with and downstream from the cooling system compressor; acooling system turbine in fluid communication with and downstream fromthe cooling system precooler; a discharge conduit downstream from thecooling system turbine; and a bypass line in fluid communication withand downstream from the cooling system precooler. The bypass line is influid communication with and upstream from the discharge conduit, andprovides selective fluid communication between an inlet side and adischarge side of the cooling system turbine to bypass the coolingsystem turbine.

Jet engines are also provided that include: an engine compressor; acombustor in flow communication with the engine compressor; an engineturbine in flow communication with the combustor to receive combustionproducts from the combustor; and a cooling system as described above influid communication with bleed air from the engine compressor. Airplanesthat include such a jet engine and at least one of an aircraft thermalmanagement system and an aircraft environmental control system, whereinthe discharge conduit is in fluid communication with the at least one ofthe aircraft thermal management system and the aircraft environmentalcontrol system.

Methods are also generally provided for cooling bleed air in a jetengine. In one embodiment, the method includes: extracting bleed airfrom a jet engine compressor; directing the bleed air to a firstprecooler, wherein the bleed air has an extracted temperature; reducingthe extracted temperature of the bleed air to a second temperature inthe first precooler; thereafter, directing a first portion of the bleedair to a heat exchanger and directing a second portion of the bleed airto a cooling system compressor such that the first portion and thesecond portion define separate flow paths; flowing the first portion ofthe bleed air through a heat exchanger to reduce the second temperatureof the first portion to a third temperature; flowing the second portionof the bleed air sequentially through a cooling system compressor, acooling system precooler, and a cooling system turbine to reduce thesecond temperature of the second portion to a fourth temperature,wherein the fourth temperature is less than the extracted temperature;and thereafter, mixing the first portion and the second portion in adischarge conduit.

These and other features, aspects and advantages of the presentinvention will become better understood with reference to the followingdescription and appended claims. The accompanying drawings, which areincorporated in and constitute a part of this specification, illustrateembodiments of the invention and, together with the description, serveto explain the principles of the invention.

BRIEF DESCRIPTION OF THE DRAWINGS

A full and enabling disclosure of the present invention, including thebest mode thereof, directed to one of ordinary skill in the art, is setforth in the specification, which makes reference to the appendedfigures, in which:

FIG. 1 is a schematic diagram of an exemplary embodiment of a cold aircooling system, shown by way of example as part of an airplane;

FIG. 2 is a schematic diagram of another exemplary embodiment of a coldair cooling system, shown by way of example as part of an airplane;

FIG. 3 is a schematic diagram of a yet another exemplary embodiment of acold air cooling system, shown by way of example as part of an airplane;

FIG. 4 is a schematic diagram of a still another exemplary embodiment ofa cold air cooling system, shown by way of example as part of anairplane;

FIG. 5 is a schematic diagram of a yet another exemplary embodiment of acold air cooling system, shown by way of example as part of an airplane;

FIG. 6 is an elevated perspective view of an airplane incorporating acold air cooling system in accordance with the instant disclosure; and

FIG. 7 illustrates a cross-sectional view of one embodiment of a gasturbine engine that may be utilized within an aircraft in accordancewith aspects of the present subject matter.

FIG. 8 shows a diagram of an exemplary method embodied by aspects of thepresent subject matter.

FIG. 9 shows a diagram of another exemplary method embodied by aspectsof the present subject matter.

DETAILED DESCRIPTION OF THE INVENTION

Reference now will be made in detail to embodiments of the invention,one or more examples of which are illustrated in the drawings. Eachexample is provided by way of explanation of the invention, notlimitation of the invention. In fact, it will be apparent to thoseskilled in the art that various modifications and variations can be madein the present invention without departing from the scope or spirit ofthe invention. For instance, features illustrated or described as partof one embodiment can be used with another embodiment to yield a stillfurther embodiment. Thus, it is intended that the present inventioncovers such modifications and variations as come within the scope of theappended claims and their equivalents.

As used herein, the terms “first”, “second”, and “third” may be usedinterchangeably to distinguish one component from another and are notintended to signify location or importance of the individual components.Also, the terms “upstream” and “downstream” refer to the relativedirection with respect to fluid flow in a fluid pathway. For example,“upstream” refers to the direction from which the fluid flows, and“downstream” refers to the direction to which the fluid flows.

A bleed air cooling system is generally provided that is in fluidcommunication with bleed air from a compressor within a jet engine. Thebleed air cooling system is generally configured to cool the receivedbleed air and provide the cooled air (e.g., via a discharge conduit fromthe bleed air cooling system) to at least one of an aircraft thermalmanagement system and an aircraft environmental control system. Thecooled air output from the cooling system has, in one embodiment, atemperature that is less than the mean flow path temperature of aircoming into and through the engine. Methods of providing cooled fluid(e.g., cooled air) are also generally provided, with the input air beingextracted from the jet engine (e.g., bleed air from the enginecompressor).

Referencing FIGS. 1 and 6, a first exemplary cold air cooling system 100is configured to supply unregulated bleed air from a high pressurecompressor (HPC) of a jet engine 112 to an aircraft thermal managementsystem and/or an environmental control system (ECS) 121 of an aircraft122 at a temperature that is below the engine air stream 110. Forpurposes of the instant disclosure, an engine cooling stream includes,without limitation, one or more of the following: fan stream air, inletair drawn into the intake, and ram air. Pursuant to the followingexemplary explanation, the system 100 of the jet engine 112 will bedescribed as being in fluid communication with structures associatedwith the aircraft 122.

Referencing FIGS. 1 and 6, a conventional thermal management system(TMS) loop 150 is utilized to draw thermal energy away from the aircraft122 and deliver/pump this thermal energy to the jet engine TMS 170. Inorder to better differentiate those structures associated with aircraft122 from structures of the engine 112, a dotted line 124 is depicted.Consequently, structures to the right of the dotted line 124 aredepicted and described in exemplary form as part of the engine 112,whereas those structures to the left of the dotted line 124 are depictedand described in exemplary form as separate from the engine 112 and areassociated with the aircraft 122. It should be understood, however, thatcertain of the components associated with the aircraft 122 could insteadbe part of the engine 112, and vice versa. Accordingly, those skilled inthe art should understand that the structures and description areexemplary in nature and the identification of structures as being partof the engine 112 or part of the aircraft 122 is not limiting.

Referring to FIG. 1, an exemplary cold air cooling system 100 is shownand includes a bleed air inlet feed 200 that is unregulated from the HPCsection of the engine 112. This bleed air inlet feed 200 supplies highpressure and high temperature compressed air to a first precooler 210.In exemplary form, this first precooler 210 facilitates the transfer ofthermal energy from the high pressure and high temperature bleed air tocooler air that is drawn into the engine 112. The bleed air output fromthe precooler 210 may have a significantly reduced temperature, but itspressure is not significantly changed. This lower temperature, highpressure air is fed into a cooling system compressor 220, whichincreases the temperature and pressure of the air. Consequently, the airoutput from the compressor 220 is significantly more pressurized andhigher in temperature than the air input to the compressor. This veryhigh pressure, high temperature air output from the compressor isdirected to a cooling system precooler 230. By way of example, thecooling system precooler 230 facilitates the transfer of thermal energyfrom the very high pressure and high temperature bleed air to flow pathtemperature air that is drawn into the engine 112. The bleed air outputfrom the cooling system precooler 230 has a significantly reducedtemperature, but its very high pressure will not be significantlychanged. This very high pressure, lower temperature bleed air outputfrom the cooling system precooler 230 is directed into a cooling systemturbine 240 having a variable area turbine nozzle (VATN). It should benoted, however, that a multiple position turbine nozzle or a fixed areaturbine nozzle may be used in lieu of the variable area turbine nozzle.Work performed by the very high pressure bleed air turning the turbine240 is utilized to power the compressor 220, with the output bleed airhaving a significantly reduced pressure and temperature. Those skilledin the art will understand that the turbine 240 may be mechanically orfluidically linked to the compressor 220 to transfer the work resultingfrom the very high pressure air expanding through the turbine.

In exemplary form, the temperature of the bleed air output from theturbine 240 into the discharge conduit 245 is lower than the mean flowpath temperature of air coming into the engine 112. This is in starkcontrast to prior art bleed air cooling systems that were unable todeliver bleed air to the aircraft 122 at a temperature below that of themean flow path temperature of air coming into the engine 112.

As shown in FIG. 2, an alternate exemplary cold air cooling system 280includes the structures of the first exemplary cold air cooling system100 and, for illustration purposes only, will be explained for use withthe TMS loop 150 from the first exemplary embodiment. Accordingly, likereference numerals refer to similar structures as discussed pursuant tothe first exemplary embodiment and will not be repeated in furtheranceof brevity.

In addition to the structures of the first exemplary cold air coolingsystem, this first alternate exemplary cold air cooling system 280includes a bypass line 290 and a control valve 292 in series with thebypass line. In exemplary form, the bypass line 290 is connected betweenthe inlet and outlet of the turbine 240 to selectively allow airdischarged from the cooling system precooler 230 to be directed to theECS 121 without traveling through the turbine. By way of example, thecontrol valve 292 is communicatively coupled to a thermocouple (notshown) in thermal communication with air discharged from the turbine240.

Depending upon the operating conditions of the engine 112 and theambient air properties (temperature, pressure, etc.), it may beadvantageous to have bleed air bypass the turbine 240. For example, ifthe temperature of the air being discharged from the turbine 240 is toolow, the control valve 292 may receive temperature readings from thethermocouple and, based upon program parameters, open or close valves inseries with the bypass line 290 in order to increase and control thetemperature of bleed air delivered to the aircraft 122 within apredetermined range. Alternatively, or in addition, the control valve292 may be in communication with a pressure sensor at the discharge ofthe turbine 240. In instances where the discharge pressure is too low,the control valve 292 may receive pressure readings from the pressuresensor and, based upon program parameters, open or close valves inseries with the bypass line 290 in order to increase and control thepressure of bleed air delivered to the aircraft 122 within apredetermined range. While the foregoing bypass has been described ashaving active management, those skilled in the art will understand thatpassive management is likewise feasible.

As shown in FIG. 3, a further alternate exemplary cold air coolingsystem 700 includes the structures of the first alternate exemplary coldair cooling system 280. Accordingly, like reference numerals refer tosimilar structures as discussed pursuant to the first alternateexemplary embodiment and will not be repeated in furtherance of brevity.

In addition to the structures of the first alternate exemplary cold aircooling system 280, this further alternate exemplary cold air coolingsystem 700 includes a bypass line 710 and a control valve 720 downstreamfrom and in fluid communication with the outlet side of the firstprecooler 210. In exemplary form, the bypass line 710 is connectedbetween the outlet of a heat exchanger 730 and the outlet of the turbine240 to selectively allow air discharged from the heat exchanger 730 tobe directed downstream from the turbine 240. The heat exchanger 730receives mean flow path air drawn into the engine 112 and uses this airas a heat sink to transfer thermal energy from the higher temperatureair exiting the first precooler 210. By way of example, the controlvalve 720 may be communicatively coupled to a thermocouple (not shown)in thermal communication with air discharged from the turbine 240.Alternatively, the control valve 720 may be passively controlled andpositioned downstream from the heat exchanger 730 in order to increasethe temperature and/or pressure of the air delivered to the aircraft122.

A discharge valve 712 is optionally positioned in the bypass line 710and in fluid communication with and downstream from the heat exchanger730, but upstream from the discharge conduit 245. The discharge valve712 is configured to control fluid flow from the heat exchanger 730 tothe discharge conduit 245. The discharge valve 712 may receivetemperature readings from the thermocouple and, based upon programparameters, open or close valves in series with the bypass line 710 inorder to increase and control the temperature of bleed air delivered tothe aircraft 122 within a predetermined range. Alternatively, or inaddition, the discharge valve 712 may be in communication with apressure sensor at the discharge of the turbine 240. In instances wherethe discharge pressure is too low, the discharge valve 712 may receivepressure readings from the pressure sensor and, based upon programparameters, open or close valves in series with the bypass line 710 inorder to increase and control the pressure of bleed air delivered to theaircraft 122 within a predetermined range. While the foregoing bypasshas been described as having active management, those skilled in the artwill understand that passive management is likewise feasible.

Depending upon the operating conditions of the engine 112 and theambient air properties (temperature, pressure, etc.), it may beadvantageous to have bleed air bypass the turbine 240. For example, ifthe temperature of the air being discharged from the turbine 240 is toolow, the control valve 720 may receive temperature readings from thethermocouple and, based upon program parameters, open or close valves inseries with the bypass line 710 in order to increase and control thetemperature of bleed air delivered to the aircraft 122 within apredetermined range. Alternatively, or in addition, the control valve720 may be in communication with a pressure sensor at the discharge ofthe turbine 240. In instances where the discharge pressure is too low,the control valve 720 may receive pressure readings from the pressuresensor and, based upon program parameters, open or close valves inseries with the bypass line 710 in order to increase and control thepressure of bleed air delivered to the aircraft 122 within apredetermined range.

Referring to FIG. 4, a second exemplary cold air cooling system 300 isconfigured to supply unregulated bleed air from a high pressurecompressor (HPC) section of a jet engine 112 to an aircraft 122 at atemperature that is below the temperature of the air flowing into theengine 112. For illustration purposes only, the second exemplary coldair cooling system 300 will be explained for use with the TMS loop 150from the first exemplary embodiment. Accordingly, like referencenumerals refer to similar structures as discussed pursuant to the firstexemplary embodiment and will not be repeated in furtherance of brevity.

As with the first exemplary embodiment, the second exemplary cold aircooling system 300 includes a bleed air inlet feed 400 that isunregulated from the HPC section of the engine 112. This bleed air inletfeed 400 supplies high pressure and high temperature compressed air to afirst precooler 410. In exemplary form, this first precooler 410facilitates the transfer of thermal energy from the high pressure andhigh temperature bleed air to air that is drawn into the engine 112. Thebleed air output from the precooler 410 may have a significantly reducedtemperature, but its pressure will not be significantly changed. Thislower temperature, high pressure air is fed into a turbine 440 having avariable area turbine nozzle. As with the foregoing embodiments, thevariable area turbine nozzle may be replaced with a multiple positionturbine nozzle or a fixed area turbine nozzle.

Work performed by the high pressure bleed air turning the turbine 440may be utilized to power other equipment associated with the engine 112or aircraft 122, with the output bleed air having a significantlyreduced pressure and temperature. By way of example, the turbine 440 maybe utilized to power a generator 475, mechanically rotate gears of agearbox 480, drive a pump 485, or any combination of the foregoingmechanical device utilized for transfer of work associated with eitherthe engine 112 or aircraft 122. Those skilled in the art will understandthat the turbine 440 may be mechanically or fluidically linked to one ormore of the foregoing components to capitalize upon the work performedby the high pressure bleed air rotating the turbine. In exemplary form,the temperature of the bleed air output from the turbine 440 is lowerthan the flow path temperature of air coming out of the precooler 410.This, again, is in stark contrast to prior art bleed air cooling systemsthat were unable to deliver bleed air to the aircraft 122 at atemperature below that of the ambient flow path air coming into theengine 112.

Referencing FIG. 5, a third exemplary cold air cooling system 500 isconfigured to supply unregulated bleed air from a high pressurecompressor (HPC) section of a jet engine 112 to an aircraft 122 at atemperature that is below the temperature of the air flowing into theengine 112. For illustration purposes only, the third exemplary cold aircooling system 500 will be explained for use with the TMS loop 150 fromthe first exemplary embodiment. Accordingly, like reference numeralsrefer to similar structures as discussed pursuant to the first exemplaryembodiment and will not be repeated in furtherance of brevity.

As with the first exemplary embodiment, the third exemplary cold aircooling system 500 includes a bleed air inlet feed 600 that isunregulated from the HPC section of the engine 112. This bleed air inletfeed 600 supplies high pressure and high temperature compressed air to afirst precooler 610. In exemplary form, this first precooler 610facilitates the transfer of thermal energy from the high pressure andhigh temperature bleed air to air that is drawn into the engine 112. Thebleed air output from the precooler 610 may have a significantly reducedtemperature, but its pressure will not be significantly changed. Thislower temperature, high pressure air is fed into a turbine 640 having avariable area turbine nozzle. As with the foregoing embodiments, amultiple position turbine nozzle or a fixed area turbine nozzle may beused in lieu of the variable area turbine nozzle.

Work performed by the high pressure bleed air turning the turbine 640may be utilized to power other equipment associated with the engine 112or aircraft 122, with the output bleed air having a significantlyreduced pressure and temperature. By way of example, the turbine 640 isutilized to power a compressor 650 on the outlet side of an enginecooler 660. By way of example, the engine cooler 660 draws in flow pathtemperature air via an inlet 670 at a predetermined pressure. The flowpath temperature air acts as a thermal sink to draw heat away from aheat source associated with the cooler 660 and exits the cooler at apressure lower than the outlet pressure discharge 680 of the compressor650. The pressure differential across the compressor 650 is operative topull air into the compressor and ultimately through the inlet 670. Thoseskilled in the art will understand that the turbine 640 may bemechanically or fluidically linked to the compressor 650 to capitalizeupon the work performed by the high pressure bleed air rotating theturbine. In exemplary form, the temperature of the bleed air output fromthe turbine 640 is lower than the flow path temperature of air comingout of the precooler 610. This, once again, is in stark contrast toprior art bleed air cooling systems that were unable to deliver bleedair to the aircraft 122 at a temperature below that of the air cominginto the engine 112. Additionally, the bleed air output (i.e., theoutput fluid) can have an output temperature that is less than abouthalf of the extracted temperature of the bleed air (e.g., less thanabout a third of the extracted temperature).

Referring to FIG. 8, an exemplary method 800 is shown for providingregulated air to an aircraft thermal management system having a thermalload. The method includes receiving and cooling at 802 from a jet engineinto a first precooler. Generally, the unregulated air has an inputpressure and an input temperature when received into the firstprecooler. Then, the first precooler can cool the unregulated air to afirst temperature. A first portion 803 of the unregulated air can thenbe compressed at 804 to a first pressure via a compressor (e.g., acooling system compressor). The first portion of the unregulated air canthen be optionally cooled at 806 (e.g., in a cooling system precooler).The first portion of unregulated air is then received at 808 into aturbine (e.g., in a cooling system turbine) having a variable areaturbine nozzle. The second portion 809 of the unregulated air is cooledat 810 via a second precooler to a second temperature. The first portionof the unregulated air and the second portion of the unregulated air arein separate flow paths, as shown. Finally, the first portion ofunregulated air in the second portion of unregulated air are regulatedat 812 to a discharge temperature and a discharge pressure selected tomeet requirements of the aircraft thermal management system byextracting work from the turbine. In certain embodiments, the extractedwork can be provided to a mechanical device utilized for transfer ofwork, such as a compressor, a gearbox, a generator, or a pump.

Referring to FIG. 9, an exemplary method 900 is shown for providingregulated air to an aircraft thermal management system having a thermalload. The method includes receiving and cooling at 902 from a jet engineinto a first precooler. Generally, the unregulated air has an inputpressure and an input temperature when received into the firstprecooler. Then, the first precooler can cool the unregulated air to afirst temperature. The unregulated air is then compressed at 904 to afirst pressure via a compressor (e.g., in a cooling system compressor).Then, the compressed unregulated air can be optionally cooled at 906(e.g., in a cooling system precooler). A first portion 907 of theunregulated air is then received and expanded at 908 into a turbinehaving a variable area turbine nozzle. The second portion 909 of theunregulated air bypasses at 910 the turbine from the cooling systemcompressor. In a discharge conduit, the first portion of unregulated airand the second portion of unregulated air are regulated at 912 to adischarge temperature and a discharge pressure selected to meetrequirements of the aircraft thermal management system by extractingwork from the turbine. In certain embodiments, the extracted work can beprovided to a mechanical device utilized for transfer of work, such as acompressor, a gearbox, a generator, or a pump.

In such methods, the first and second precoolers can use fan stream airfrom the jet engine as a heat sink fluid, with the fan stream air havinga fan stream temperature and a fan stream pressure. In one embodiment,the discharge temperature is less than the fan stream temperature. Forexample, the jet engine can operate at sea level static conditions witha fan stream pressure that is above about 17 psi at idle and/or aboveabout 30 psi at take-off.

The regulated air can provide a reduction of more than about 10% of thethermal load of the aircraft thermal management system, such as morethan about 60%. For example, the regulated air can provide reduction ofmore than about 4 kW of thermal load to about 90 kW of thermal load.

It should be noted, however, that air other than flow path air may beutilized as the thermal sink for any of the precoolers 210, 230, 410,610, 730. Moreover, while the foregoing exemplary embodiments have beendescribed as including precoolers, it should be understood that aprecooler is synonymous with a heat exchanger.

It should also be understood that while the systems 100, 280, 300, 500,700 have been described as being associated with a jet engine 112, it isalso within the scope of the disclosure to have these systems in fluidcommunication with other vehicles for use on water or land (e.g., boatsand automobiles).

FIG. 7 illustrates a cross-sectional view of one embodiment of a gasturbine engine 112 that may be utilized within an aircraft in accordancewith aspects of the present subject matter, with the engine 112 beingshown having a longitudinal or axial centerline axis 12 extendingtherethrough for reference purposes. Although shown as a turbofan jetengine, any suitable jet engine can be utilized with the cooling systemdescribed herein. For example, suitable jet engines include but are notlimited to high-bypass turbofan engines, low-bypass turbofan engines,turbojet engines, turboprop engines, turboshaft engines, propfanengines, and so forth.

As shown in FIG. 7, the exemplary engine 112 may include a core gasturbine engine (indicated generally by reference character 14) and a fansection 16 positioned upstream thereof. The core engine 14 may generallyinclude a substantially tubular outer casing 18 that defines an annularinlet 20. In addition, the outer casing 18 may further enclose andsupport a booster compressor 22 for increasing the pressure of the airthat enters the core engine 14 to a first pressure level. A highpressure, multi-stage, axial-flow compressor 24 may then receive thepressurized air from the booster compressor 22 and further increase thepressure of such air. The pressurized air exiting the high-pressurecompressor 24 may then flow to a combustor 26 within which fuel isinjected into the flow of pressurized air, with the resulting mixturebeing combusted within the combustor 26. The high energy combustionproducts are directed from the combustor 26 along the hot gas path ofthe engine 10 to a first (high pressure) turbine 28 for driving the highpressure compressor 24 via a first (high pressure) drive shaft 30, andthen to a second (low pressure) turbine 32 for driving the boostercompressor 22 and fan section 16 via a second (low pressure) drive shaft34 that is generally coaxial with first drive shaft 30. After drivingeach of turbines 28 and 32, the combustion products may be expelled fromthe core engine 14 via an exhaust nozzle 36 to provide propulsive jetthrust.

Additionally, as shown in FIG. 7, the fan section 16 of the engine 10may generally include a rotatable, axial-flow fan rotor 38 thatconfigured to be surrounded by an annular fan casing 40. It should beappreciated by those of ordinary skill in the art that the fan casing 40may be configured to be supported relative to the core engine 14 by aplurality of substantially radially-extending, circumferentially-spacedoutlet guide vanes 42. As such, the fan casing 40 may enclose the fanrotor 38 and its corresponding fan rotor blades 44. Moreover, adownstream section 46 of the fan casing 40 may extend over an outerportion of the core engine 14 so as to define a secondary, or by-pass,airflow conduit 48 that provides additional propulsive jet thrust.

During operation of the engine 10, it should be appreciated that aninitial air flow (indicated by arrow 50) may enter the engine 10 throughan associated inlet 52 of the fan casing 40. The air flow 50 then passesthrough the fan blades 44 and splits into a first compressed air flow(indicated by arrow 54) that moves through conduit 48 and a secondcompressed air flow (indicated by arrow 56) which enters the boostercompressor 22. The pressure of the second compressed air flow 56 is thenincreased and enters the high pressure compressor 24 (as indicated byarrow 58). After mixing with fuel and being combusted within thecombustor 26, the combustion products 60 exit the combustor 26 and flowthrough the first turbine 28. Thereafter, the combustion products 60flow through the second turbine 32 and exit the exhaust nozzle 36 toprovide thrust for the engine 10.

As used herein, flow path temperature refers to the median temperatureof a fluid, namely air, while it is flowing through a predeterminedsection/stage of a jet engine. More specifically, flow path temperaturecan be no lower than the lowest temperature air drawn into the intake ofa jet engine. As air is drawn into the intake and compressed, the airwill increase in temperature and, accordingly, have a flow pathtemperature near the end of the compressor section that is above that ofthe flow path temperature of air at the intake.

It should also be noted that while the foregoing exemplary embodimentshave been described in the context of an aircraft, the instantdisclosure is equally applicable to vehicles beyond aircraft. Anyvehicle having cabin or other cooling needs may be addressed using thepresent disclosure. For example, a boat that is jet powered may benefitfrom the exemplary embodiments of the instant disclosure to providecooling to the cabin and/or the electronics associated with the boat.Consequently, the foregoing disclosure is by no means limited inapplication to aircraft, but rather is applicable to any vehicleutilizing jet power to provide a cooling stream wherever beneficial.Those skilled in the art will readily recognize the utility of thepresent disclosure in the context of other vehicles.

Following from the above description, it should be apparent to those ofordinary skill in the art that, while the methods and apparatuses hereindescribed constitute exemplary embodiments of the present disclosure, itis to be understood that the disclosures contained herein are notlimited to the above precise embodiments and that changes may be madewithout departing from the scope of the disclosure. Likewise, it is tobe understood that it is not necessary to meet any or all of theidentified advantages or objects of the disclosure in order to fallwithin the scope of the disclosure, since inherent and/or unforeseenadvantages of the present disclosure may exist even though they may nothave been explicitly discussed herein.

What is claimed is:
 1. A cooling system in fluid communication withbleed air from a jet engine compressor, the cooling system comprising: afirst precooler in fluid communication with an engine air stream and influid communication with the bleed air from the jet engine compressor; acooling system compressor in fluid communication, via the bleed air,with and downstream from the first precooler; a cooling system precoolerin fluid communication with the engine air stream upstream of the firstprecooler, and in fluid communication, via the bleed air, with anddownstream from the cooling system compressor; a cooling system turbinein fluid communication, via the bleed air, with and downstream from thecooling system precooler; a discharge conduit in fluid communication,via the bleed air, with and downstream from the cooling system turbine;a bypass line in fluid communication, via the bleed air, with anddownstream from the cooling system precooler, wherein the bypass line isin fluid communication, via the bleed air, with and upstream from thedischarge conduit, and wherein the bypass line provides selective fluidcommunication between an inlet side and a discharge side of the coolingsystem turbine to bypass the cooling system turbine; and a heatexchanger in fluid communication with the engine air stream upstream ofthe first precooler and downstream of the cooling system precooler, andin fluid communication fluid, via the bleed air, with and downstreamfrom the first precooler, the heat exchanger in fluid communication, viathe bleed air, with and upstream from the cooling system compressor, andthe heat exchanger and the cooling system compressor in separate bleedair flow paths from the first precooler.
 2. The cooling system of claim1, wherein the cooling system turbine has a variable area turbinenozzle.
 3. The cooling system of claim 2, wherein the variable areaturbine nozzle is configured to control fluid flow from the coolingsystem turbine to the discharge conduit.
 4. The cooling system of claim1, further comprising: a control valve in series with the bypass line.5. The cooling system of claim 4, wherein the control valve isconfigured to control fluid flow through the bypass line to bypass thecooling system turbine.
 6. The cooling system of claim 1, wherein thebleed air has an extracted temperature, and wherein the dischargeconduit provides an output fluid having an output temperature that isless than about half of the extracted temperature of the bleed air. 7.The cooling system of claim 1, wherein the engine air stream to thefirst precooler and the cooling system precooler is fan stream air,inlet air, ram air, or a combination thereof from the jet engine used asa heat sink fluid, the heat sink fluid having an engine air streamtemperature and engine air stream pressure, and wherein the dischargeconduit provides an output fluid having an output temperature that isless than the engine air stream temperature.
 8. The cooling system ofclaim 1, further comprising: a control valve in fluid communication withand downstream from the first precooler, wherein the control valvecomprises a first outlet in fluid communication with and upstream fromthe heat exchanger, and wherein the control valve comprises a secondoutlet in fluid communication with and upstream from the cooling systemturbine, and further wherein the control valve is configured to regulatecontrol fluid flow to each of the heat exchanger and the cooling systemturbine.
 9. A method of cooling bleed air in a jet engine, the methodcomprising: providing a cooling system comprising a first precooler influid communication with an engine air stream and in fluid communicationwith the bleed air from a jet engine compressor, a cooling systemcompressor in fluid communication, via the bleed air, with anddownstream from the first precooler, a cooling system precooler in fluidcommunication with the engine air stream upstream of the first precoolerand in fluid communication, via the bleed air, with and downstream fromthe cooling system compressor, a cooling system turbine in fluidcommunication, via the bleed air, with and downstream from the coolingsystem precooler, a discharge conduit in fluid communication, via thebleed air, with and downstream from the cooling system turbine, a bypassline in fluid communication, via the bleed air, with and downstream fromthe cooling system precooler, wherein the bypass line is in fluidcommunication in fluid communication, via the bleed air, with andupstream from the discharge conduit, and wherein the bypass lineprovides selective fluid communication between an inlet side and adischarge side of the cooling system turbine to bypass the coolingsystem turbine, and a heat exchanger in fluid communication with theengine air stream upstream of the first precooler and downstream of thecooling system precooler, and in fluid communication, via the bleed air,with and downstream from the first precooler, the heat exchanger influid communication, via the bleed air, with and upstream from thecooling system compressor, and the heat exchanger and the cooling systemcompressor in separate flow paths from the first precooler; extractingthe bleed air from the jet engine compressor; directing the bleed air tothe first precooler, wherein the bleed air has an extracted temperature;reducing the extracted temperature of the bleed air to a secondtemperature in the first precooler; thereafter, directing a firstportion of the bleed air to the cooling system compressor; flowing thefirst portion of the bleed air sequentially through the cooling systemcompressor and the cooling system precooler; flowing the first portionof the bleed air from the cooling system precooler into the coolingsystem turbine to reduce the second temperature of the first portion ofthe bleed air to a third temperature, wherein the third temperature isless than the second temperature; flowing a second portion of the bleedair from the cooling system precooler through the bypass line; andthereafter, mixing the first portion and the second portion in thedischarge conduit.
 10. The method as in claim 9, further comprising:controlling the flow of the first portion through the cooling systemturbine.
 11. The method as in claim 10, wherein the cooling systemturbine has a variable area turbine nozzle to control the flow throughthe cooling system turbine.
 12. The method as in claim 9, furthercomprising: controlling the flow of the first portion through thecooling system turbine.
 13. The method as in claim 12, wherein a controlvalve is in series with the bypass line.
 14. The method as in claim 13,wherein the control valve is configured to control a flow of the secondportion through the bypass line to bypass the cooling system turbine.15. The method as in claim 9, wherein the bleed air has an extractedtemperature, and wherein the discharge conduit provides an output fluidhaving an output temperature that is less than about half of theextracted temperature of the bleed air.
 16. The method as in claim 9,wherein the engine air stream to the first precooler and the coolingsystem precooler is fan stream air, inlet air, ram air, or a combinationthereof from the jet engine used as a heat sink fluid, the heat sinkfluid having an engine air stream temperature and engine air streampressure, and wherein the discharge conduit provides an output fluidhaving an output temperature that is less than the engine air streamtemperature.
 17. A method of providing regulated air to an aircraftthermal management system having a thermal load; the method comprising:providing a cooling system comprising a first precooler in fluidcommunication with an engine air stream and in fluid communication witha bleed air from a jet engine compressor, a cooling system compressor influid communication, via the bleed air, with and downstream from thefirst precooler, a cooling system precooler in fluid communication withthe engine air stream upstream of the first precooler, and in fluidcommunication, via the bleed air, with and downstream from the coolingsystem compressor, a cooling system turbine in fluid communication, viathe bleed air, with and downstream from the cooling system precooler, adischarge conduit downstream from the cooling system turbine, a bypassline in fluid communication, via the bleed air, with and downstream fromthe cooling system precooler, wherein the bypass line is in fluidcommunication, via the bleed air, with and upstream from the dischargeconduit, and wherein the bypass line provides selective fluidcommunication between an inlet side and a discharge side of the coolingsystem turbine to bypass the cooling system turbine, and a heatexchanger in fluid communication with the engine air stream upstream ofthe first precooler and downstream of the cooling system precooler, andin fluid communication, via the bleed air, with and downstream from thefirst precooler, the heat exchanger in fluid communication, via thebleed air, with and upstream from the cooling system compressor, and theheat exchanger and the cooling system compressor in separate flow pathsfrom the first precooler; receiving unregulated air from a jet engineinto the first precooler, the unregulated air having an input pressureand an input temperature; cooling the unregulated air in the firstprecooler to a first temperature; compressing the unregulated air viathe cooling system compressor to a first pressure; cooling theunregulated air via the cooling system precooler to a second pressure;receiving a first portion of the unregulated air downstream from thecooling system compressor into the cooling system turbine having avariable area turbine nozzle; expanding the first portion of theunregulated air in the cooling system turbine and into the dischargeconduit; bypassing a second portion of the unregulated air downstreamfrom the cooling system compressor into the discharge conduit; andexpanding the first portion of unregulated air to a dischargetemperature and a discharge pressure selected to meet requirements ofthe aircraft thermal management system by extracting work from thecooling system turbine.
 18. The method of claim 17, wherein the engineair stream to the first precooler and the cooling system precooler isfan stream air, inlet air, ram air, or a combination thereof from thejet engine used as a heat sink fluid, the heat sink fluid having anengine air stream temperature and engine air stream pressure.
 19. Themethod of claim 18, wherein the discharge temperature is less than theengine air stream temperature.
 20. The method of claim 17, wherein theextracted work is provided to a mechanical device utilized for transferof work, the device selected from the group consisting of a compressor,a gearbox, a generator, and a pump.
 21. The method of claim 17, whereinthe regulated air provides a thermal load reduction of 4 kW to 90 kW.